Compact lightweight ramjet engines incorporating swirl augmented combustion with improved performance

ABSTRACT

A ramjet powered device that utilizes a novel swirl generator for rapidly and efficiently atomizing, vaporizing, as necessary, and mixing a fuel into an oxidant. The swirl generator converts an oxidant flow into a turbulent, three-dimensional flowfield into which the fuel is introduced. The swirl generator effects a toroidal outer recirculation zone and an inner central recirculation zone, both of which are configured in a backward-flowing manner that carries heat and combustion byproducts upstream where they are employed to continuously ignite a combustible fuel/oxidizer mixture in adjacent shear layers and stabilizes flame propagation and accelerates combustion throughout the entire combustor. The swirl generator provides smooth combustion with no instabilities and minimum total pressure losses and enables significant reductions in the L/D ratio of the combustor. Other benefits include simplicity, reliability, wide flammability limits and high combustion efficiency and thrust performance.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a division of U.S. application Ser. No. 10/360,469, filed Feb. 6, 2003 entitled “Compact Lightweight Ramjet Engines Incorporating Swirl Augmented Combustion With Improved Performance”, now U.S. Pat. No. 6,968,695, which is a continuation-in-part of U.S. application Ser. No. 10/243,961, filed Sept. 13, 2002 entitled “Compact, Lightweight High-Performance Lift Thruster Incorporating Swirl-Augmented Oxidizer/Fuel Injection Mixing and Combustion”, which issued as U.S. Pat. No. 6,820,411 on Nov. 23, 2004. Other features of the present invention are discussed and claimed in U.S. application Ser. No. 10/360,168, now U.S. Pat. No. 6,907,724, entitled “Combined Cycle Engines Incorporating Swirl Augmented Ramjet Combustion for Reduced Volume and Weight and Improved Performance” and in U.S. application Ser. No. 10/360,016, now U.S. Pat. No. 6,895,756 entitled “Compact Swirl Augmented Afterburners for Gas Turbine Engines”.

FIELD OF THE INVENTION

The present invention generally relates to ramjet engines and more particularly to the improvement of fuel/oxidizer mixing systems and ramjet engines.

BACKGROUND OF THE INVENTION

The performance of a ramjet powered device, such as a missile, is dependent on the size, weight and performance of the individual components of the missile as well as the interaction between these components when integrated into an overall “system”. In general, if the performance of an individual component is increased and/or the size and/or weight of the individual component is reduced, the overall performance of the ramjet powered device will be improved.

The component that has the most impact on the design and performance of the ramjet engine and the device it powers is the combustor. As those skilled in the art will appreciate, traditionally combustor designs that feature short combustor lengths typically require high blockage instream flameholders in the combustor that cause large pressure losses and therefore reduce propulsion performance. In contrast, long combustors without the relatively large instream flameholders typically increase the overall size and weight of a ramjet powered device and as such, effectively reduce the range and/or payload capacity of the ramjet powered device.

Classically designed ramjet powered missiles are known to employ combustors with instream bluffbody flameholders, which resulted in a combustor length-to-diameter (L/D) ratio of about 3.0. To accelerate these classically designed ramjets to their supersonic operational speeds a strapped-on external rocket booster was used. Due to the external booster these designs were relatively bulky.

The relative bulk and high weight of classically designed ramjet missiles prompted development of more streamlined designs that were thought to be more cost effective and have longer ranges and/or greater payload capacity. One such design, the integral rocket ramjet (IRR), employs an integral rocket booster, housed inside the ramjet combustor, for accelerating the missile to its ramjet takeover speed, which is typically about Mach 1.5 or higher. With the combined, dual use, rocket/ramjet combustor the liquid fueled IRR is typically sized with a combustor L/D ratio of about 5.0 creating a relatively long and heavy propulsion system.

A combination of the classical and streamlined designs resolved some of the problems with the prior ramjet powered devices, but the combustor required an L/D ratio of about 3.0, was still relatively long and too heavy. Furthermore, significant pressure losses were associated with the instream flameholders and as such, thrust performance and mission capability were degraded. Accordingly, there remains a need in the art for a ramjet engine having an improved combustor.

SUMMARY OF THE INVENTION

In one form, the present teachings provide an apparatus having at least one core turbojet engine, at least one ramjet engine and a flow controller. Each core turbojet engine has a low pressure compressor, a high pressure compressor, a combustor, and a turbine. Each ramjet engine is coupled to one of the core turbojet engines and has an air inlet, a swirl generator, a ramjet combustor and a nozzle. The swirl generator is coupled to the air inlet and is operable for converting an oxidizer flow into a three-dimensional flowfield that includes a tangential velocity component. The swirl generator including a flow defining means, which is operable for effecting both an outer recirculation zone and a central recirculation zone in the ramjet combustor, and a fueling means. The outer recirculation zone is toroidal in shape while the central recirculation zone is disposed inwardly of the outer recirculation zone. The fueling means is operable for fueling the outer and central recirculation zones. Heat and combustion by-products produced during combustion in the ramjet combustor are carried upstream by the outer and central recirculation zones where the heat and combustion by-products are employed to continuously ignite a combustible fuel/oxidizer mixture in a shear layer adjacent each of the outer and central recirculation zones, which accelerate propagation of the flame into and throughout the core flowfield. The flow controller is coupled to the core turbojet engines and the ramjet engines and has at least one forward moveable element and at least one aft moveable element. The at least one forward moveable element is movable between a first position, in which the air intake of the ramjet engines is closed and atmospheric air is permitted to flow into the low pressure compressor of the core turbojet engines, and a second position in which atmospheric air is not permitted to flow into the lower pressure compressor of the core turbojet engines and the air intake of the ramjet engines is open. The at least one aft moveable element is movable between a first position, in which an exit of the nozzle is closed and an exit of the core turbojet engines is open to permit exhaust gas to flow therethrough, and a second position in which the exit of the nozzle is open and the exit of the core turbojet engines is closed.

In another form, the teachings of the present invention provide a method that can include: providing a device having a core turbojet engine, a ramjet engine and a flow controller, the core turbojet engine having a low pressure compressor, a high pressure compressor, a combustor, and a turbine, the ramjet engine being coupled to the core turbojet engine and having an air inlet, a swirl generator, a ramjet combustor and a nozzle, the swirl generator being coupled to the air inlet and being operable for converting an oxidizer flow into a three-dimensional flowfield that includes a tangential velocity component, the swirl generator including a flow defining means and a fueling means, the flow defining means being operable for effecting both an outer recirculation zone and a central recirculation zone in the ramjet combustor, the outer recirculation zone being toroidal in shape, the central recirculation zone being disposed inwardly of the outer recirculation zone, the fueling means being operable for fueling the outer and central recirculation zones, wherein heat and combustion by-products produced during combustion in the ramjet combustor are carried upstream by the outer and central recirculation zones where the heat and combustion by-products are employed to continuously ignite a combustible fuel/oxidizer mixture in a shear layer adjacent each of the outer and central recirculation zones, which accelerate propagation of the flame into and throughout the core flowfield, the flow controller being coupled to the core turbojet engine and the ramjet engine, the flow controller having at least one forward moveable element and at least one aft moveable element, the at least one forward moveable element being movable between a first position, in which the air intake of the ramjet engine is closed and atmospheric air is permitted to flow into the low pressure compressor of the core turbojet engine, and a second position in which atmospheric air is not permitted to flow into the lower pressure compressor of the core turbojet engine and the air intake of the ramjet engine is open, the at least one aft moveable element being movable between a first position, in which an exit of the nozzle is closed and an exit of the core turbojet engine is open to permit exhaust gas to flow therethrough, and a second position in which the exit of the nozzle is open and the exit of the core turbojet engine is closed; positioning the forward and aft movable elements in the first position if a velocity of the device is not greater than a first transition velocity and operating only the core turbojet engine; and operating both the core turbojet engine and the ramjet engine if the velocity of the device is greater than the first transition velocity and less than a second transition velocity.

The present invention overcomes the aforementioned drawbacks by providing a ramjet engine having a swirl generator that causes extremely rapid fuel atomization, vaporization and mixing with a combustion efficiency that ranges from about 90% to about 99% enabling the use of combustor lengths that are substantially shorter than that of the prior art.

Further areas of applicability of the present invention will become apparent from the detailed description provided hereinafter. It should be understood that the detailed description and specific examples, while indicating the preferred embodiment of the invention, are intended for purposes of illustration only and are not intended to limit the scope of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

Additional advantages and features of the present invention will become apparent from the subsequent description and the appended claims, taken in conjunction with the accompanying drawings, wherein:

FIG. 1 is a perspective view of a jet aircraft having a pair of lift thrust augmentors constructed in accordance with the teachings of the present invention;

FIG. 2 is a schematic illustration of a portion of the jet aircraft of FIG. 1;

FIG. 3 is a cut-away perspective view of a portion of the lift thrust augmentor illustrated in FIG. 1;

FIG. 4 is a longitudinal section view of a portion of the lift thrust augmentor;

FIG. 5 is an exploded perspective view of a portion of the lift thrust augmentor illustrating the elbow in greater detail;

FIG. 6 is a longitudinal section view of the elbow;

FIG. 7 is an exploded perspective view of a portion of the lift thrust augmentor;

FIG. 8 is an exploded perspective view of a portion of the lift thrust augmentor illustrating the centerbody hub assembly in greater detail;

FIG. 9 is a longitudinal section view of the centerbody hub assembly;

FIG. 9A is a schematic view of the swirl generator of the present invention illustrating several various centerbody assembly and wall injector fueling schemes;

FIG. 10 is a perspective view of a portion of the swirl generator illustrating the swirl vane pack in greater detail;

FIG. 10A is a partial top plan view of an alternately constructed swirl vane pack illustrating a profiled vane configuration;

FIG. 11 is an exploded perspective view of the swirl vane pack of FIG. 10;

FIG. 12 is an exploded sectional view of a portion of the swirl vane pack;

FIG. 13 is a longitudinal section view similar to that of FIG. 4 but illustrating a combustion event;

FIG. 14 is a partially broken away perspective view of a portion of a swirl generator constructed in accordance with the teachings of an alternate embodiment of the present invention;

FIG. 15 is a sectional view of a portion of a swirl generator constructed in accordance with the teachings of a second alternate embodiment of the present invention which illustrates an alternatively constructed centerbody hub assembly in detail;

FIG. 16 is an exploded perspective view of a portion of a swirl generator constructed in accordance with the teachings of a third alternate embodiment of the present invention which illustrates another alternatively constructed centerbody hub assembly;

FIG. 17 is a longitudinal section view of the centerbody hub assembly of FIG. 16;

FIG. 18 is a partially broken away side elevation view of a swirl generator constructed in accordance with the teachings of a fourth alternate embodiment of the present invention;

FIG. 19 is a perspective view of a portion of a swirl generator constructed in accordance with the teachings of a fifth alternative embodiment of the present invention which illustrates a vane for an alternative swirl vane pack with a plurality of fuel injection sites;

FIG. 20 is a partial sectional view of the vane of FIG. 19;

FIG. 21 is a perspective view of a portion of a swirl generator constructed in accordance with the teachings of a sixth alternative embodiment of the present invention which illustrates a vane for an alternative swirl vane pack with turbulator ramps and a plurality of fuel injection sites;

FIG. 22 is a partial sectional view taken through the vane of FIG. 21;

FIG. 23 is a front elevation view of a swirl generator constructed in accordance with the teachings of a seventh alternative embodiment of the present invention which illustrates a vane for a third alternative swirl vane pack having scallops;

FIG. 24 is a partial sectional view of a swirl generator constructed in accordance with the teachings of a eighth alternative embodiment of the present invention which illustrates an alternative inlet housing wherein a plurality of channels are formed into the inlet ramp;

FIG. 24A is a sectional view taken along the line 24A—24A of FIG. 24;

FIG. 25 is a partial sectional view of a swirl generator constructed in accordance with the teachings of a ninth alternative embodiment of the present invention which illustrates another alternative inlet housing illustrating the incorporation of fuel injection sites into the inlet ramp;

FIG. 26 is a partial sectional view of a swirl generator constructed in accordance with the teachings of a tenth alternative embodiment of the present invention which illustrates an alternative combustor illustrating the use of a quarl extension;

FIG. 27 is a partial sectional view of a swirl generator constructed in accordance with the teachings of a eleventh alternative embodiment of the present invention which illustrates another alternately configured centerbody hub assembly;

FIG. 28 is a partial sectional view of a swirl generator constructed in accordance with the teachings of a twelfth alternative embodiment of the present invention which illustrates another alternately configured centerbody hub assembly;

FIG. 29 is a front elevation view of a ramjet missile that incorporates a swirl generator constructed in accordance with the teachings of the present invention;

FIG. 30 is a longitudinal section view taken along the line 30—30 of FIG. 29;

FIG. 30A is a view similar to FIG. 30 but illustrating a swirl generator that employs flow guide vanes;

FIG. 31 is a longitudinal section view of a ramshell that incorporates a swirl generator constructed in accordance with the teachings of the present invention;

FIG. 32 is a partial longitudinal section view of a combined cycle engine having a plurality of ramjet engines that incorporate a swirl generator constructed in accordance with the teachings of the present invention;

FIG. 33 is a section view taken along the line 33—33 of FIG. 32;

FIG. 34 is a partial longitudinal section view of a second combined cycle engine having a plurality of ramjet engines that incorporate a swirl generator constructed in accordance with the teachings of the present invention;

FIG. 35 is a section view taken along the line 35—35 of FIG. 34;

FIG. 35A is a partial longitudinal section view similar of a third combined cycle engine having a ramjet engine that incorporates a swirl generator constructed in accordance with the teachings of the present invention;

FIG. 36 is a longitudinal section view of a rocket-based combined cycle engine constructed in accordance with the teachings of the present invention;

FIG. 37 is a rear elevation view of the rocket-based combined cycle engine of FIG. 37;

FIG. 38 is a perspective view of a portion of the rocket-based combined cycle engine illustrating the variable area throat in an open condition;

FIG. 39 is a perspective view similar to that of FIG. 38, but illustrating the variable area throat in a closed condition;

FIG. 40 is a longitudinal section view of an aircraft engine having a conventional afterburner;

FIG. 41 is a longitudinal section view similar to FIG. 40 but illustrating an afterburner that incorporates a swirl generator constructed in accordance with the teachings of the present invention; and

FIG. 41A is a longitudinal section view similar to that of FIG. 41 but illustrating an afterburner that utilizes a swirl generator with a collapisable centerbody cone.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

With reference to FIG. 1 of the drawings, an exemplary jet aircraft 8 is illustrated to include a pair of lift thrust augmentors 10 that are constructed in accordance with the teachings of the present invention. A conventional gas turbine engine 12 serves as the primary source of propulsive power for the jet aircraft 8, while the lift thrust augmentors 10 are selectively operable to produce thrust when the demand for thrust exceeds a predetermined threshold.

With additional reference to FIGS. 2 through 4, each lift thrust augmentor 10 includes an air delivery portion 20, a swirl generator 30 and a combustor/nozzle portion 40. When combined, the swirl generator 30 and combustor/nozzle portion 40 are also known as a compact swirl-augmented thruster (COSAT). The air delivery portion 20 includes a butterfly valve 50 and in the particular example provided, an elbow 60. It will be readily apparent to those skilled in the art that the air delivery portion 20 may be configured in any manner to package the lift thrust augmentor 10 into a particular application and as such, the elbow 60 may be omitted or the bend angle changed to suit the specific needs of a given application.

The butterfly valve 50 and the elbow 60 are coupled in fluid connection to the gas turbine engine 12. More specifically, hot high-pressure air is bled from the bypass fan 12 a of the gas turbine engine 12 and diverted through a lateral jetscreen feed system (not specifically shown) to an attitude control system 70. The butterfly valve 50, which is normally maintained in a closed condition and coupled in fluid connection with the attitude control system 70, is opened to divert a predetermined amount (i.e., mass flow rate) of the hot high-pressure bleed air to the lift thrust augmentors 10. In the particular embodiment provided, about 30% of the airflow that is directed into the attitude control system 70 is redirected to the lift thrust augmentors 10 when the lift thrust augmentors 10 are operated to provide maximum thrust.

The air that passes by the butterfly valve 50 is routed through the elbow 60 to the swirl generator 30. As will be described in greater detail below, the swirl generator 30 injects fuel in to the air and promotes the efficient mixing of the air and fuel. A torch igniter 80, such as that described in copending U.S. patent application Ser. No. 10/217,972 entitled “Torch Igniter”, the disclosure of which is hereby incorporated by reference as if fully set forth herein, is employed to initiate a combustion event wherein the fuel/air mixture is burned in the combustor 40 a of the combustor/nozzle portion 40. Those skilled in the art will appreciate that igniters, including electric spark igniters, plasma jet igniters, lasers and microwaves, may be employed in the alternative. As the construction of the combustor 40 a is well known in the art, a detailed discussion of the combustor will not be provided herein. For example, those skilled in the art will appreciate that the combustor 40 a may be wholly formed of a suitable high temperature material, or may utilize a perforated liner that facilitates air-cooling of the combustor 40 a, or may utilize a wall that is partially or wholly comprised of fluid conduits that facilitate a flow of fluid through or about the combustor wall that operates to cool the combustor 40 a during its operation.

Thereafter, hot combustion by-products are expelled through a nozzle 40 b, such as a convergent nozzle (FIG. 4) or a convergent/divergent nozzle (FIG. 2), to produce thrust. It will be understood that the particular combustible mixture (i.e., a liquid fuel and air) that is utilized in this example is not intended to limit the scope of the disclosure in any way. In this regard, those skilled in the art will understand that any type of fuel (e.g., liquid, slurry, gas) and any type of oxidant (e.g., air, hydrogen peroxide, oxygen) may be utilized in the swirl generator and lift thrust augmentor of the present invention. Accordingly, when the term “mixing” is used in the context of a fuel/oxidant mixture, it will be understood to include atomization and vaporization if the fuel is not injected in gaseous state.

With reference to FIGS. 5 and 6, the elbow 60 is illustrated to turn the airflow 90 about 90° and is relatively constant in its cross-sectional area. The elbow 60 includes an elbow housing 94, a plurality of flow guide vanes 96, a pair of side covers 98 and an aerodynamic fairing 100. As will be readily apparent to one skilled in the art, the components of the elbow 60 may be fabricated out of any appropriate material, the selection of which is largely dependant on the temperature of the air entering the air delivery portion 20 (FIG. 2). In the particular example provided, the inlet temperature may be relatively high and as such, materials including Cress stainless steel series (e.g., 304, 321), Haynes 240, ceramic matrix composites (e.g., C/SiC, SiC/SiC, Si₃N₄, Al₂O₃/Al₂O₃) would be appropriate for the construction of the components of the elbow 60. As those skilled in the art will appreciate, various thermally protective coatings (e.g., ceramics) and/or heat transfer techniques that rely on a cooling effect provided by a working fluid (e.g., fuel, air) may additionally or alternatively be employed to render the components of the elbow 60 suitable for a given set of inlet conditions.

The elbow housing 94 is shown to include an inlet portion 102 and an outlet portion 104 that are interconnected by an arcuate turning portion 106. The elbow housing 94 may be unitarily formed via conventional processes such as casting or machining (from bar stock), or may be a multi-component fabrication that is fixedly coupled together as by welding. The lateral sidewalls 110 of the turning portion 106 include a plurality of concentric slots 112 that are formed therethrough, while the endwall 114 in the turning portion 106 includes a fairing aperture 116. The fairing aperture 116 is sized to receive a portion of the aerodynamic fairing 100, as well as a conduit 120 (FIG. 4) that couples the swirl generator 30 to an ignition controller 122 (FIG. 4) that actuates the lift thrust augmentor 10, as well as the fuel source 126 (FIG. 4) of the jet aircraft 8.

The flow guide vanes 96 are curved about a single axis and slide into the turning portion 106 through the concentric slots 112. Accordingly, the flow guide vanes 96 extend across the inner dimension of the turning portion 106 in a direction that is generally perpendicular to the direction of the air flow and concentric with the radius of the turn in the turning portion 106. The flow guide vanes 96 are also configured such that they extend between the inlet portion 102 and the outlet portion 104 of the elbow housing 94. Accordingly, the radially inward flow guide vanes, such as flow guide vane 96 a, are relatively shorter than the radially outer flow guide vanes, such as flow guide vane 96 d.

The leading and trailing edges 130 and 132, respectively, of the flow guide vanes 96 preferably engage the opposite ends of the concentric slots 112, while the opposite lateral sides 136 of the flow guide vanes 96 are abutted against the side covers 98, which are fixedly secured to the elbow housing 94 via a plurality of screws (not specifically shown). Additionally or alternatively, the flow guide vanes 96 may be welded in place to secure them in place within the concentric slots 112.

The flow guide vanes 96 are employed to mitigate air flow distortions and flow separation induced by the associated upstream butterfly valve 50 (FIG. 2) as well as to prevent further flow separation, the creation of secondary flows and large scale profile distortions due to centrifugal forces as the inlet flow undergoes the 90° turn through the turning portion 106. If unabated, secondary flows, separations and distortions tend to complicate the design and operation of the downstream fuel injection, alter the mixing effectiveness and provide stimuli for combustion instabilities. Consequently, the configurations of the air delivery portion 20 (generally) and the elbow 60 (specifically) provide a high degree of uniformity in the flow of air (i.e., a uniform axial airflow) to the swirl generator 30.

As best seen in FIG. 6, the flow guide vanes 96 cooperate with the elbow housing 94 to define a plurality of inlet flow channels 140. As the path that is defined by inlet flow channel 140 a is relatively shorter between the inlet portion 102 and the outlet portion 104 as compared with the inlet flow channel 140 e and as transient flow differences between the inlet flow channels 140 are highly undesirable, the flow guide vanes 96 of the particular embodiment illustrated are positioned in a radial direction in a manner that provides inlet flow channels 140 with dissimilar cross-sectional areas such that the inlet flow channels 140 produce a series of inlet flows that are relatively uniform in flow velocity. Accordingly, the inlet flow 150 to the swirl generator 30 has a velocity that is substantially completely defined by an axial velocity component.

In FIGS. 4 through 6, the conduit 120 that couples the swirl generator 30 to the ignition controller 122 (FIG. 2) and the fuel source 126 (FIG. 2) is illustrated to extend through the endwall 114 of the turning portion 106, through a conduit aperture 154 in the flow guide vane 96 d and out the outlet portion 104 of the elbow housing 94 where it is coupled to the swirl generator 30. The presence of the conduit 120 in the interior of the turning portion 106 would ordinarily initiate a flow separation, which as mentioned above, reduces the effectiveness and efficiency of the lift thrust augmentor 10. The aerodynamic fairing 100, however, is employed to reduce or eliminate altogether, the flow separations that would be induced by the presence of the conduit 120.

In the example provided, the aerodynamic fairing 100 includes a hollow hub 156 and an airfoil portion 158 that is coupled to the hub 156. The conduit 120 extends through the hollow interior of the hub 156 and may be coupled thereto by any appropriate retaining means, including an interference fit (e.g., shrink fit or press fit), brazing and welding. The airfoil portion 158, which surrounds the conduit 120, extends in the in the down-wind direction in a tapered manner and terminates at its trailing edge 164 with a relatively small amount of trailing edge bluntness. Alternatively, the airfoil portion 158 may be configured to abut the windward side of the conduit 120. The airfoil portion 158 is positioned within the flow channel 140 e so as to reduce or eliminate altogether the flow separations that would be induced by the presence of the conduit 120. The hub 156 and/or airfoil portion 158 may be fixedly coupled to the elbow housing 94 by any appropriate retaining means, including an interference fit (e.g., shrink fit or press fit), brazing and welding.

With reference to FIGS. 4 and 7, the swirl generator 30 is illustrated to include an inlet housing 200, a centerbody hub assembly 204, a swirl vane pack 206 and a wall injector assembly 208. In the particular example provided, the inlet housing 200 is illustrated to include an upper inlet housing segment 200 a, which is coupled to and integrally formed with the outlet portion 104 of the elbow 60, an optional wall injection housing 228 (which will be described in detail, below) and an optional lower inlet housing segment 200 b, which is coupled to the combustor/nozzle portion 40 and which includes an inlet extension 210 and an inlet ramp 212. Alternatively, the inlet housing 200 may be unitarily formed. Also alternatively, the upper inlet housing segment 200 a may be separate from the outlet portion 104 of the elbow 60 and/or the lower inlet housing segment 200 b, if included, may be integrally formed with the combustor/nozzle portion 40. Accordingly, those skilled in the art will appreciate that one or more of the upper and lower inlet housing segments 200 a and 200 b and the wall injection housing 208 may not exist as a discrete component. In the presently preferred configuration of the lift thrust augmentor 10, the upper inlet housing segment 200 a is integrally formed with the elbow 60 so as to minimize the overall length of the lift thrust augmentor 10.

The inlet housing 200 defines a hollow interior volume 220 into which the centerbody hub assembly 204 extends. The injection of fuel into the lift thrust augmentor 10 is illustrated to include fuel injection through the wall 224 of the inlet housing 200 at a location forwardly of the inlet ramp 212. The sizing and purpose of the inlet ramp 212 will be discussed in greater detail, below.

As all fuel injection occurs downstream of the swirl vane pack 206, the presence of the inlet extension 210 effectively shifts the various points of fuel injection in an upstream direction relative to the inlet ramp 212 so as to provide additional time for the fuel to mix (i.e., for the liquid fuel of this example to atomize, mix and vaporize) prior to entering the combustor 40 a. Those skilled in the art will appreciate that the amount of additional mixing time that is afforded by the inlet extension 210 is a function of its length. Those skilled in the art will also appreciate that the need for an inlet extension is based on the state of the air flow, such as the velocity, temperature, pressure and the characteristics of the fuel that is being used. Therefore, in some of the applications that we have conceived for the swirl generator 30, such as ramjets, the air flow is at a sufficiently high temperature such that even liquid fuels are rapidly mixed (i.e., atomized, mixed and vaporized), which permits the aft end of the centerbody hub assembly 204 to be positioned so as to extend downstream of the inlet ramp 212 into the combustor 40 a.

In the particular example provided, wall injection is accomplished through the wall injector assembly 208. The wall injector assembly 208 includes an annular wall injection housing 228 (which is considered to be part of the injection housing 200) and a plurality of fuel injectors 230 that are circumferentially spaced about the wall injection housing 228. The inside diameter of the wall injection housing 228 is equal in diameter to the inner diameter of the lower inlet housing segment 200 b so as to be flush and not to induce flow separations, etc. that would tend to impede the efficiency of the swirl generator 30. The inside diameter wall injection housing 228, however, is somewhat smaller in diameter than the inside diameter of the upper inlet housing segment 200 a for reasons that will be described in detail, below.

The wall injection housing 228 is disposed between the aft end of the upper inlet housing segment 200 a and the forward end of the lower inlet housing segment 200 b. Conventional Viton O-rings 238 or other sealing devices that are well known in the art are employed to seal the interface between the opposite faces of the wall injection housing 228 and the upper and lower inlet housing segments 200 a and 200 b. Threaded fasteners (not shown) are employed to fixedly couple the elbow 60, the upper inlet housing segment 200 a, the wall injection housing 228, the lower inlet housing segment 200 b and the combustor/nozzle portion 40 together.

In the example shown, the fuel injectors 230 comprise eight flush-mount simplex fuel injectors such as Woodward FST Simplex Injectors, which are commercially available from Woodward FST, Inc. of Zeeland, Mich. Those skilled in the art will appreciate, however, that other amounts and/or types of fuel injectors, including stand-off or wall flush simplex fuel injectors, orifice injectors or variable area poppet fuel injectors with variations in drop sizes and cone angles (i.e., solid or hollow cone) may also be used.

In FIGS. 7 through 9, the centerbody hub assembly 204 is illustrated to include the conduit 120 and a centerbody assembly 236, which includes a forward centerbody 240, a conduit retainer 242, and an aft centerbody 244 and an igniter 246. In the particular embodiment illustrated, the igniter 246 is a conventional spark igniter which is commercially available from sources such as those that are manufactured by Champion Spark Plug Company of Toledo, Ohio. Accordingly, a detailed description of the construction of the igniter 246 will not be provided herein. Those skilled in the art will appreciate that other types of igniters, such as a plasma jet igniter (also well known in the art and commercially available from sources such as Unison Industries Inc. of Jacksonville Fla.), microwave ignition devices, and laser ignition devices may be employed as an alternative to a spark igniter.

As noted above, the conduit 120 extends through the elbow 60 (FIG. 4) to couple the swirl generator 30 (FIG. 4) to the ignition controller 122 (FIG. 4) and the fuel source 126 (FIG. 4). In the particular embodiment provided, the conduit 120 is a hollow tube into which an electrical cable 250 and a plurality of fuel conduits 252 are housed. The electrical cable 250 electrically couples the ignition controller 122 to the igniter 246 such that electrical energy is transmitted to the igniter 246 when the ignition controller 122 is actuated to permit the igniter 246 to produce a discharge arc (not shown). The fuel conduits 252 couple in fluid connection of the fuel source 126 to a plurality of fuel injectors in the aft centerbody 244. Those skilled in the art will appreciate, however, that a single fuel conduit (not shown) may be alternatively employed, wherein the single fuel conduit supplies fuel to a fuel manifold within or coupled to the aft centerbody 244 to which are coupled the fuel injectors. Also alternatively, the swirl generator 30 may be constructed without a discrete fuel conduit 252 wherein such function would additionally be provided by the conduit 120 that houses the electrical cable 250. Those skilled in the art will appreciate that these fuel splits could alternatively be accomplished through a set of orifices that effectively limit the mass flow rate of fuel therethrough.

The forward centerbody 240 is a generally hollow structure having a central aperture 270 into which the conduit 120 is fixedly coupled, as through brazing for example. The exterior surface of the forward centerbody 240 includes an aerodynamically contoured leading surface 272 as well as a mounting flange 274 to which the swirl vane pack 206 is mounted. The aerodynamically contoured leading surface 272, which is illustrated to be generally spherically shaped in the particular example provided, serves to guide the inlet flow 150 (FIG. 4) exiting the elbow 60 radially outwardly around the forward centerbody 240 and into the swirl vane pack 206. The mounting flange 274 has a diameter that is generally smaller than the trailing edge of the aerodynamically contoured leading surface 272 to thereby create an abutting flange 278 at the intersection of the mounting flange 274 and the leading surface 272. The aft end of the interior of the forward centerbody 240 includes a counterbored portion 280 that is configured to receive the conduit retainer 242.

The conduit retainer 242 is an annular structure with an open center that is configured to receive therethrough portions of the igniter 246 and the electrical cable 250. A plurality of conduit mounting apertures 284 are formed through the conduit retainer 242 and are sized to matingly receive an associated one of the fuel conduits 252. The fuel conduits 252 are preferably fixedly coupled to the conduit retainer 242 in their associated conduit mounting apertures 284 through an appropriate joining process, such as brazing. Conventional threaded fasteners, such as socket head cap screws 290, are preferably employed to fixedly but releasably couple the conduit retainer 242 to the forward centerbody 240.

The aft centerbody 244 may be configured in several different manners to optimize the efficiency of the lift thrust augmentor 10 and tailor its thrust output to a desired thrust output level. In the particular example provided the aft centerbody 244 includes first and second injector rings 300 and 302, respectfully, and an aft bluff boat-tail 304.

A conventional Viton O-ring 310 or other well known sealing device is employed to create a seal between the first injector ring 300 and the aft face of the forward centerbody 240. The first injector ring 300 has a generally hollow center through which the igniter 246 is received, and an outside diameter that is relatively larger in diameter than that of the mounting flange 274. The first injector ring 300 is illustrated to include a plurality of circumferentially spaced apart fuel injectors 320, such as simplex injectors having a flow number (FN) equal to about 8.5, which are commercially available from Woodward FST, Inc. of Zeeland, Mich. Those skilled in the art, however, will appreciate that other types of fuel injectors, including a plurality of orifices, could be additionally or alternatively employed. Although the injectors 320 are illustrated as being configured to inject fuel in a radially outward direction, those skilled in the art will appreciate that the fuel injectors 320 may be alternatively configured to inject fuel in an upstream direction, a downstream direction or any combination of the radially outward, downstream and upstream directions.

The fuel injectors 320 are coupled in fluid connection to a manifold 324 that is formed into the first injector ring 300. In the example provided, the manifold 324 is coupled in fluid connection to an associated one of the fuel conduits 252 to receive fuel therefrom. Those skilled in the art will appreciate, however, that the manifold 324 could also be coupled in fluid connection to each of the fuel conduits 252. A Viton O-ring 326 or other well known sealing device is employed to seal the interface between the conduit retainer 242 and the front face of the first injector ring 300.

A conventional Viton O-ring 330 or other well known sealing device is employed to create a seal between the second injector ring 302 and the aft face of the first injector ring 300. The second injector ring 302 has a generally hollow center that is at least partially threaded so as to threadably engage a threaded portion 336 of the igniter 246 in a conventional manner. The second injector ring 302 may include no fuel injectors (i.e., constitute a “blank” injector ring as shown in FIG. 9A), or may include a plurality of circumferentially spaced fuel injectors 340, such as simplex or orifice injectors, depending on the desired output of the lift thrust augmentor 10. Although the injectors 340 are illustrated as being configured to inject fuel in a radially outward direction (FIG. 8), those skilled in the art will appreciate that the fuel injectors 340 may be alternatively configured to inject fuel in an upstream direction, a downstream direction or any combination of the radially outward, downstream and upstream directions.

The second injector ring 302 may be keyed or otherwise aligned to the first injector ring 300 in any conventional manner to maintain the first and second injector rings 300 and 302 in a condition wherein they are aligned about a common axis. In this condition, the outer surface of the second injector ring 302 substantially coincides with the outer surface of the first injector ring 300 to thereby prevent the generation of any flow separations or discontinuities.

In the embodiments wherein the second injector ring 302 includes fuel injectors 340, a manifold 344, which is formed into the second injector ring 302, is employed to couple in fluid connection the fuel injectors 340 to an associated fuel conduit 252. The manifold 344 may be coupled in fluid connection to the manifold 324 of the first injector ring 300, or to an aperture 345 that is formed through the first injector ring 300 as is shown in FIG. 8. The interface between the first and second injector rings 300 and 302 is sealed by a Viton O-ring 346 in an area proximate the aperture 345.

The aft bluff boat-tail 304 includes a flange portion 360 and a flow-effecting portion 362. The flange portion 360 abuts the aft face of the second injector ring 302 and includes a pair of apertures 364 through which conventional socket head cap screws 366 are received. The cap screws 366 extend through similar apertures formed in the first and second injector rings 300 and 302 and the conduit retainer 242 and threadably engage apertures (not shown) in the forward centerbody 240 to fixedly couple these components to one another. The flange portion 360 also includes a hollow center 370 into which a tip 246 a of the igniter 246 extends. The hollow center 370 is chamfered on its aft end so as to provide additional space about the tip 246 a for a flame kernel.

The flow-effecting portion 362 is coupled to the outer perimeter of the flange portion 360. In the particular embodiment provided, the flow-effecting portion 362 is frusto-conically shaped and includes a plurality of circumferentially spaced apart fuel injectors 380 that are configured to inject fuel in a predetermined direction. Like the second injector ring 302, the flow-effecting portion 362 may be alternately configured to include no fuel injectors (i.e., constituting a “blank” bluff body). Although the injectors 380 are illustrated as being configured to inject fuel in an upstream direction, those skilled in the art will appreciate that the fuel injectors 380 may be alternatively configured to inject fuel in a radially outward direction, a downstream direction or any combination of the radially outward, downstream and upstream directions.

In the example illustrated, the flow-effecting portion has an initial outer diameter that matches the outer diameter of the first and second injector rings 300 and 302. The flow-effecting portion 362 terminates at its aft end at a sharp edge 386 that operates to initiate flow separation and to anchor and radially extend the central recirculation zone 610 (FIG. 13) to increase its size and flameholding capabilities.

Although the centerbody hub assembly 204 has been illustrated thus far as being formed from a plurality of discrete components, those skilled in the art will appreciate that various known manufacturing techniques, including direct metal fabrication, may be employed so as to reduce the actual number of components that are utilized. For example, the forward centerbody 240, the swirl vane pack 206 and the first injector ring 300 may all be unitarily formed, which would thereby eliminate the need for the conduit retainer 242.

Referring to FIG. 9A, the swirl generator 30 of the present invention is schematically illustrated to show several of the various fueling options that may be employed. The fuel injectors 230 may comprise, for example, orifices A, flush-mount simplex injectors B or stand-off simplex injectors C. The fuel injectors of the aft centerbody assembly 244 (i.e., fuel injectors 320, 340, 380 and 840) may include simplex injectors D or orifices E, or may be omitted in part (blank) as designated by reference letter F.

In FIGS. 4, and 10 through 12, the swirl vane pack 206 is illustrated to include a mounting hub 400, a plurality of vanes 402 and a shroud 404. In the particular example provided, the swirl vane pack 206 is an assembly wherein the components comprising the swirl vane pack 206 are fabricated, assembled and fixedly secured to one another. Those skilled in the art will appreciate, however, that alternative fabrication techniques may be employed to reduce the number of components that comprise the swirl vane pack 206. For example, technologies such as hot isostatic pressing, casting and direct metal fabrication may be employed to form the swirl vane pack 206, either wholly or partially, or in combination with the centerbody assembly 236 or portions thereof as described above.

The mounting hub 400 is an annular structure that is received over the mounting flange 274 (FIG. 9) of the forward centerbody 240 (FIG. 9) in juxtaposed relation with the abutting flange 278 (FIG. 9) of the forward centerbody 240 and the front face of the first injector ring 300 (FIG. 9). The cap screws 366 (FIG. 8) exert a clamping force that fixedly but removably secures the aft centerbody 244 to the forward centerbody 240. As the mounting hub 400 abuts the mounting flange 274 and the front face of the first injector ring 300, the clamping force is also transmitted between the abutting flange 278 and the front face of the first injector ring 300, which operates to fixedly secure (both axially and radially) the mounting hub 400 therebetween.

The vanes 402 of the swirl vane pack 206 are configured with a swirl number that ranges from about 0.4 to about 1.2 so as to permit the combustor/nozzle portion 40 to achieve a combustor length-to-diameter (L/D) ratio (as measured from a plane at which the dump step 636 in FIG. 13 to the throat of the nozzle portion 40 b) that is less than about 2.0, and preferably less than about 1.6 and more preferably about 1.0 or less. In the particular example provided, we utilized twelve vanes 402 having a straight configuration that is skewed to the centerline of the mounting hub 400 so as to provide a swirl number of 0.54. Those skilled in the art will appreciate that various other vane configurations may alternatively be employed, including vanes with different skew angles and/or an arcuate or helical profile (FIG. 10A). As the swirl vane pack 206 is comprised of a plurality of discrete components, the vanes 402 are configured with tabs 410 a and 410 b on their opposite ends. The radially inward tabs 410 a are configured to engage apertures 412 that are formed on the opposite faces of the mounting hub 400. The tabs 410 a and the apertures 412 cooperate to align the vanes 402 to the mounting hub 400 so that the vanes 402 may be coupled to the mounting hub 400 in a conventional manner, such as brazing or welding.

The shroud 404 includes a pair of end caps 420 and a pair of circumferentially extending portions 422. The end caps 420 include a plurality of apertures 424 that are configured to receive the radially outward tabs 410 b on the vanes 402. In a manner similar to that of the apertures 412 of the mounting flange 274, the apertures 424 cooperate to align the vanes 402 to the end caps 420. The circumferentially extending portions 422 are disposed around the perimeter of the vanes 402 between the tabs 410 b and the end caps 420, circumferentially extending portions 422 and vanes 402 are fixedly secured together, as through welding.

With reference to FIG. 4, the swirl vane pack 206 is illustrated to be fixedly coupled to the centerbody hub assembly 204 in the manner described above and disposed between the wall injector assembly 208 and a recessed step 500 formed in the upper inlet housing segment 200 a. In this location, the trailing edges of the vanes 402 of the swirl vane pack 206 are located upstream of all fuel injection sites, which eliminates any potential for flashback which would damage the vanes 402. The shroud 404 of the swirl vane pack 206 is preferably sized to engage in a press-fit manner the recessed step 500 in the lower inlet housing segment 200 b to thereby structurally couple the swirl vane pack 206 and the centerbody hub assembly 204 to the inlet housing 200.

The swirl generator 30 is configured such that the vanes 402 impart tangential velocities to the axial inlet flow 150 to convert the inlet flow 150 into a spiraling, three dimensional swirling flow structure or flowfield 510 (FIG. 4). The flowfield 510 has a dramatic effect on the rate of fuel mixing, atomization, droplet vaporization, flame propagation, combustion efficiency, combustion stability, combustion intensity and widens flammability limits. Those whom are skilled in the art will appreciate that radial velocities will be affected by the swirling effect, but that the major impact of swirling effect concerns the aforementioned tangential velocity component.

With reference to FIG. 13, the high tangential velocities produced by the vanes 402, whether straight or profiled in their configuration, creates a very intense shear layer 600 and enhances the large scale vortex or central recirculation zone 610 that is generated and anchored by the bluff end 614 of the aft centerbody 244, even at relatively low levels of swirl (i.e., a level of swirl that is greater than or equal to about 0.4). In the case of vanes 402 having a flat configuration, especially high shear stresses are created that promote very efficient mixing of the fuel that is introduced into the inlet housing 200 via the wall injector assembly 208 and the injectors 320, 340 and 380 in the aft centerbody 244 due to high intensity turbulence that is generated by the trailing edge vortices that are induced by the flow separation on the lee side of the vanes 402. Combustion in the central recirculation zone 610 is initiated by a flame kernel 630 that is produced by the igniter 246 that is housed in the centerbody hub assembly 204. Additionally or alternatively, igniters 80 a (similar or identical to igniter 246 or igniter 80) may also be employed near the dump plane 636 a.

The inlet ramp 212, which is optional, aids in increasing the size of the dump step 636 that occurs at the dump plane 636 a. In the particular example provided, the inlet ramp 212 helps to create a relatively large 90° dump step 636 at the transition between the inlet housing 200 and the inlet of the combustor 40 a that serves to considerably improve flame propagation rates and the combustor's operability limits. More specifically, the dump step 636 creates a toroidal outer recirculation zone 640 along the combustor wall that is initially ignited by a flame kernel that is produced by the torch igniter 80 or the igniter(s) 246 and/or 80 a (FIG. 4). The length of the outer recirculation zone 640 is a function of the height of the dump step 636 and the strength of the swirl number of the swirl vane pack 206. Generally speaking, for a given constant swirl number, the length, size and robustness of the outer recirculation zone 640 are directly related to the height of the dump step. The inlet ramp 212 and its shape not only provide a means to easily tune the flow height of the dump step 636 and the flow direction, but also increases the local flow velocities to thereby intensify the separated shear layer turbulence and increases the rate of mass entrainment of fuel into the shear layer. The ramp shape and reduced flow gap height also accelerates early merging of the shear layers 650 and 600 of the outer recirculation zone 640 and the central recirculation zone 610, respectively, which is essential for combustors having a relatively short length. Those skilled in the art will appreciate that the shape of the ramp can be altered to change the maximum height and, therefore, the volume of the outer recirculation zone and the gaps between the outer and central recirculation zones.

Portions of the fuel that are dispensed by the fuel injectors 320, 340 and 380 are employed to substantially fuel the central recirculation zone 610, while portions of the fuel that are dispensed by the fuel injectors 230 of the wall injector assembly 208 are employed to fuel both the outer recirculation zone 640 and fuel the central recirculation zone 610. Any portions of the fuel that is dispensed by the injectors 230, 320, 340 and 380 that is not employed to fuel the central or outer recirculation zones 610 and 640 is employed to generally fuel the main core flow 700, which as those skilled in the art will appreciate, consists of the entire flow of combusting fuel and air except the central recirculation zone 610 and the outer recirculation zone 640. The central recirculation zone 610 and the outer recirculation zone 640, once formed, contain a fixed trapped mass. An exchange of mass occurs between each of the central and outer recirculation zones 610 and 640 and the core flow 700, but there is no net change in mass for steady flow conditions in either of the central and outer recirculation zones 610 and 640.

As each of the injectors 320, 340 and 380 are coupled in a discrete manner to the fuel source 126, the amount of fuel that is dispensed by the injectors 320, 340 and 380 may be tailored in a desired manner to fine tune flame stabilization and combustion performance during throttling. Accordingly, the injectors 320, 340 and 380 may be independently controlled so as to provide a relatively wide range of flexibility to control combustor characteristics, depending on a particular application.

The shear layers 650 and 600 of the outer recirculation zone 640 and the central recirculation zone 610 provide reduced velocity regions to hold the flame, and maintain and propagate the combustion process. More specifically, the outer recirculation zone 640 and the central recirculation zone 610 provide flame stabilization and act as a robust ignition source for the core flow 700 by supplying heat and recirculated chemically reacting by-products, such as OH, H and O radicals, to the main fuel/air mixture of the combustor 40 a. In this regard, each recirculation zone carries the heat and chemically active species from the flame in the respective shear layer and recirculating flow volume upstream where they act to ignite the fresh combustible fuel/air mixture entering the shear layer to thereby provide a continuous pilot for the core flow 700.

As noted above, the aft centerbody 244 of the swirl generator 30 may be configured in various different arrangements to achieve desired design parameters. For example, the aft centerbody 244 may be configured with a channeled aft bluff boat-tail 304 a as illustrated in FIG. 14. The aft bluff boat-tail 304 a is generally similar to the aft bluff boat-tail 304 of FIG. 8, except that it includes a plurality of channels 800 that are formed about the perimeter of the flow-effecting portion 362 a. The channels 800 are formed at an angle relative to the centerline of the aft bluff boat-tail 304 a that maintains the effective flow direction provided by the swirl vane pack 206. Those skilled in the art will appreciate that other channels may be selected to control turbulent transport and mixing including ones that are opposite of the tangential direction provided by the swirl vane pack. The ramp-like geometry of the channels produces a spectrum of turbulence scales that enhances mixing to promote flame intensity and propagation from the central recirculation zone 610 into the core flow 700 (FIG. 13). Like the aft bluff boat-tail 304, the aft bluff boat-tail 304 a may include one or more fuel injectors, or may have a “blank” configuration (i.e., a configuration without one or more fuel injectors).

In the embodiment of FIG. 15, the aft centerbody 244 b is illustrated to be generally similar to the aft centerbody 244 of FIG. 9, except that it does not include a discrete bluff body. In this regard, the second injector ring 302 essentially forms a bluff body as the aft centerbody 244 b terminates abruptly at the rear face of the second injector ring 302.

The embodiment of FIGS. 16 and 17 is generally similar to that of FIG. 15 except that the igniter 246 has been replaced with a fuel injector 840, such as a simplex atomizer having, for example, a 100° spray angle. The fuel injector 840, like the previously discussed fuel injectors that are housed in the aft centerbody 244, is individually coupled to a fuel conduit 252 so as to permit the fuel injector 840 to be selectively deployed. In this embodiment, the central recirculation zone is also ignited by the flame kernel that is produced by the torch igniter 80 and/or the igniter 80 a that are described above as initiating combustion in the outer recirculation zone.

Another embodiment is illustrated in FIG. 18 wherein the fuel injectors in the wall of the inlet housing 200 f are replaced by a plurality of cross-flow strut injectors 900. Each of the cross-flow strut injectors 900 is swaged into the aft centerbody 244 f for structural support and coupled in fluid connection to fuel conduits 906 that extend through the inlet housing 200 f. Each cross-flow strut injectors 900 has a plurality of orifices 910 that promote atomization of the fuel flowing therethrough. Additionally, this embodiment includes a channeled aft bluff boat-tail 304 f and a center-mount fuel injector 840.

The embodiment of FIGS. 19 and 20 provides a fuel injection scheme that is very similar to that of the embodiment of FIG. 18. Instead of cross-flow strut injectors, however, this embodiment utilizes a plurality of fuel injection sites 1000 a that are formed into a trailing edge 1002 g of at least a portion of the vanes 402 g of the swirl vane pack. The injection sites 1000 a are coupled in fluid connection to an associated fuel conduit that extends through the conduit 120 (FIG. 4); a plurality of internal channels 1004 in the vanes 402 g serve to transmit the fuel through the vanes 402 g to the injection sites 1000 a. As illustrated in FIG. 20, each of the injection sites 1000 a is an orifice 1006 that has an appropriate length to diameter ratio and which is formed into the trailing edge 1002 g at a predetermined angle relative to an axis of the vane 402 g. Although the injection sites 1000 a are illustrated as being generally oriented in a downstream direction, those skilled in the art will appreciate that additionally or alternatively the injection sites could also be skewed to the axis of the vane 402 g. Additionally or alternatively, the holes 1006 may be formed on a lateral surface of the vanes 402 g to inject fuel in a desired direction (see, e.g., injection sites 1000 b and 1000 c in phantom in FIG. 20, which are also formed with an appropriate orifice length to diameter ratio). Despite the complexity of the vane arrangement, this embodiment is advantageous in that the exterior surfaces of the vanes 402 g form an effective means for transferring heat from the air flow to the fuel in the vanes 402 g which operates to cool the swirl vane pack as well as to increase the temperature of the fuel that is injected, which tends to increase the rate by which the fuel is mixed (i.e., improve atomization, decrease the size of the droplets in the fuel spray and directly increase the rate of droplet vaporization).

The embodiment of FIGS. 21 and 22 is generally similar to that of FIGS. 19 and 20, except that an array or row of turbulator ramps 1050 are formed or mounted onto at least a portion of the vanes 402 h of the swirl vane pack 206 to further enhance mixing. The turbulator ramps 1050 and the channels 1052 that are formed between each pair of turbulator ramps 1050 are employed to generate vortices that emanate from the trailing edge 1002 h of the vanes 402 h; the vortices enhance turbulent transport and provide highly controlled fine scale mixing. Like the previous embodiment, a plurality of fuel injection sites 1000 h that are formed into a trailing edge 1002 h (i.e., into the tubulators 1050) which are coupled in fluid connection to an associated fuel conduit that extends through the conduit 120 (FIG. 4); a plurality of integrally formed channels 1002 h in each of the vanes 402 h serve to transmit the fuel through the vanes 402 h. Those skilled in the art will also appreciate that the turbulators 1050 may also be utilized in vane configurations that do not inject fuel (i.e., in vanes without injection sites formed therein).

The example of FIG. 23 is generally similar to that of FIG. 4, except that the trailing edge 1002 i of the vanes 402 i includes a plurality of scallops 1100 rather than turbulators. The scallops 1100 are illustrated to be formed in a uniform manner wherein the crest 1102 is relatively wider than the root 1104. Those skilled in the art will appreciate, however, that the scope of the present invention is not limited to any particular scallop pattern. Those skilled in the art will appreciate, too, that the vanes 402 i may also be configured with a plurality of fuel injection sites in the manner described above for the embodiments of FIGS. 19 through 22.

In the embodiment of FIGS. 24 and 24A, the inlet ramp 212 j is illustrated to be generally similar to the inlet ramp 212 of FIG. 4, except that a plurality of channels 1150 are formed about the perimeter of the inlet ramp 212 j. The channels 1150 are formed at a predetermined angle relative to the longitudinal axis of the inlet housing 200 j and serve to enhance turbulent transport and fine scale mixing.

In the embodiment of FIG. 25, a plurality of circumferentially spaced apart fuel injection sites 1200 are formed about the circumference of the inlet ramp 212 k. The fuel injection sites 1200 are operable for injecting fuel into the flow field which aids in flame stabilization and communication between the outer and central recirculation zones. As those skilled in the art will appreciate, the channels of the previous embodiment may additionally be incorporated into the inlet ramp 212 k.

Those skilled in the art will appreciate that the wall of the combustor 40 a in an area proximate to the outer recirculation zone tends to absorb a relatively large amount of heat during the combustion process. An optional quarl expansion 1250 may be provided as shown in FIG. 26 at the dump step 636. The quarl expansion 1250 is an annular element having a generally triangular cross section; the quarl expansion 1250 is employed to “fill” the backwards facing step at the dump plane such that the angle of the dump step 636 is reduced from 90°. The angle β of the quarl expansion 1250 may be varied in a known manner to affect the size of the central recirculation zone and the rate of heat transfer to the wall of the combustor 40 a and the inlet housing 200.

The embodiment of FIG. 27 illustrates that the configuration of the centerbody assembly aft of the leading surface need not have the configuration of a right cylinder. In the particular embodiment illustrated, the portion of the centerbody assembly 236 m aft of the leading surface 272 m has a generally frusto-conical shape that is symmetrical about a longitudinal axis of the swirl generator 30 m. Those skilled in the art should also appreciate that the portion of the centerbody assembly that is positioned aft of the leading surface may be other than conical (e.g., ogival). Those skilled in the art will also recognize that multiple fuel injection sites, similar to fuel injectors 320, 340, 380 and/or 840 of FIGS. 8 and 16 could be incorporated into the centerbody assembly 236 m.

The embodiment of FIG. 28 is generally similar to the embodiment of FIG. 14 in that it includes two injection rings 1300 and 1302 that are positioned forwardly of the aft bluff boat-tail 304 a. This embodiment differs from the embodiment of FIG. 14 in that the injection rings 1300 and 1302 are configured in a manner that is generally similar to that of the first injection ring 300 and a third injection ring 1304 is coupled to the aft end of the aft bluff boat-tail 304 a. The injection ring 1304 is configured generally similar to that of the second injection ring 302 discussed above (i.e., includes fuel injectors and threadably engages the threaded portion 336 of the igniter 246). Additionally, the injection ring 1304 includes a plurality of optional grooves or channels 1310 that may be continuous with the grooves or channels of the bluff boat-tail body 800 which produces a spectrum of turbulent scales that enhance fuel and air transport and fine scale mixing to promote flame intensity and propagation from the central recirculation zone 610 into the core flow 700. As those skilled in the art will appreciate, the injection rings 1300, 1302 and/or 1304 may be “blank” and/or may include grooves or channels 1310 for producing fine scale turbulence. Furthermore, the injection ring 1304 may be configured with an aft facing fuel injector such as that which is shown in FIGS. 16 and 17.

While the swirl generator of the present invention has been illustrated and described thus far as being a component of a lift thrust augmentor, those skilled in the art will appreciate that the invention, in its broader aspects, may be utilized in diverse other applications. In FIGS. 29 and 30, for example, the swirl generator 30 is illustrated in conjunction with a ramjet missile 2000.

In this example, the ramjet missile 2000 includes a forebody 2002, fins 2004, a booster engine 2006 and a ramjet engine 2010 having an air inlet 2012, the swirl generator 30 and a ramjet combustor/nozzle 2014. The forebody 2002 conventionally houses the payload (not shown), the fuel (not shown), batteries (not shown) and the control portion (not shown) of the ramjet missile 2000, while the fins 2004 conventionally stabilize and guide the ramjet missile 2000. The air inlet 2012 includes a movable or consumable port cover (not shown) that is selectively operable for sealing the air inlet 2012 and the swirl generator during the operation of the rocket booster engine 2006. In the example provided, the rocket booster engine 2006 includes a solid propellant 2018 that burns during a boost phase of the missile's operation; hot combustion by-products are expelled from the nozzle 2020 of the booster engine 2006 to generate thrust.

Subsequent to the boost phase of the missile's operation, the booster engine 2006 of the exemplary ramjet missile illustrated is ejected and the movable port cover is also ejected or consumed so as to permit air to flow into the air inlet 2012 so that the task of propulsion may switch from the rocket booster engine 2006 to the ramjet engine 2010. The speed of the missile 2000 and the configuration of the air inlet 2012 cooperate to induce an airflow 2026 through the air inlet 2012 that is directed toward the swirl generator 30. Alternatively, as illustrated in FIG. 30A, a plurality of flow guide vanes 96 a (similar to the flow guide vanes 96 in the elbow 60) may be employed in the air inlet 2012 to provide an axisymmetric (i.e., uniform and axial) airflow to the swirl generator 30. Those skilled in the art will recognize that a chin inlet may be used followed by an annular air transfer duct that acts as an isolator and carries the air aft to the swirl generator 30 and the combustor of the ramjet combustor/nozzle 2014. In this case the air transfer duct will typically be S-shaped and may require guide vanes to straighten the flow as previously described. The swirl generator 30 is employed to generate a turbulent flowfield and to inject fuel therein in the manner described above. As in the example of the lift thrust augmentor, the swirl generator 30 operates to effect both an outer recirculation zone (proximate the dump step 2030 at the transition between the inlet housing 2032 and the inlet of the combustor/nozzle 2014), as well as a central recirculation zone (which is anchored by the aft end of the aft centerbody assembly 244).

Integration of the swirl generator 30 into the ramjet missile 2000 permits significant reductions to the L/D ratio of the combustor 2014 a and substantial reductions in the overall length and weight of the ramjet missile 2000. Furthermore, the shortened combustor length provided by the swirl generator 30 allows the booster engine 2006 to be separately packaged and thus be ejectable as the propulsion switches from the booster engine 2006 to the ramjet engine 2010. A shorter and lighter weight missile, due to the swirl generator 30, will offer significant maneuverability advantages over longer and/or heavier missiles due to a shorter turning radius capability. The COSAT technology may also require a smaller booster due to ramjet take-over occurring at a lower Mach number and hence offer additional range capability and/or further improved agility.

Another application is illustrated in FIG. 31 wherein the swirl generator 30 n is illustrated in conjunction with a ramshell 2300. The ramshell 2300 is a gun-launched, spin-stabilized projectile that employs the above-described ramjet technology to accelerate the ramshell 2300 to a velocity of about Mach 4 to about Mach 6 to extend the projectile range, minimize the time to target and maximize the penetration capability of the projectile. The ramshell 2300 includes a housing 2302, a projectile 2304, a plurality of inlet struts 2306, the swirl generator 30 n and a combustor/nozzle portion 2308.

The housing 2302 is a hollow shell that is inwardly tapered at it's front end to define an air inlet 2310. The inlet struts 2306 are fixedly coupled to the interior of the housing 2302 and to the projectile 2304 to centrally mount the projectile 2304 in a forward portion of the housing 2302. As the ramshell 2300 is spin-stabilized, the inlet struts 2306 have a spiral shape to maintain alignment with the incoming air flow. The air flow is compressed through the inlet section 2310 a and shocked to subsonic velocities near the aft end of the inlet struts 2306. While the projectile 2304 is illustrated as being a solid metallic rod, those skilled in the art will appreciate that any form of payload, including an explosive charge, may be employed in the alternative.

The swirl generator 30 n is mounted on the aft end of the projectile 2304 and the swirl vane pack 206 serves to support the housing 2302 aft of the inlet struts 2306. Fuel injection is somewhat different from that of the swirl generator 30 of FIG. 4 in that the primary purposes of the swirl generator 30 n are to augment the central recirculation zone (similar to the central recirculation zone 610 of FIG. 13) and to control the rate of mixing in the shear layer above the central recirculation zone and to control the rate of flame propagation into the core flow. Accordingly, wall injection is not employed in this embodiment, and the fuel injectors 2320 in the centerbody assembly 236 n are fueled by a reservoir 2330 that is internal to the aft centerbody assembly 244 n. The reservoir 2330 includes a pressurized bladder 2332 that surrounds the fuel 2334 to maintain the pressure of the fuel 2334 at sufficient levels during the operation of the ramshell 2300. The centrifugal force of the rotating fuel 2334 in the reservoir 2330 keeps the cooler, high density liquid fuel against the outer perimeter of the reservoir 2330 to assist in thermal protection of the combustor/nozzle portion 2308.

Fuel injection in the ramshell 2300 is preferably designed to maintain a somewhat fuel-rich condition in the central recirculation zone. As the main propulsive combustion initiates in the shear layer above the central recirculation zone, the level of mixing and heat release is controlled through the design of the swirl generator 30 n so that only the air flow in the vicinity of the shear layer participates in combustion. Operation in this manner leaves the outer region near the interior side of the housing 2302 relatively cooler and protects the housing 2302 from the high heat flux near the throat of the combustor/nozzle portion 2308, before mixing of the hot combusting gases with the outermost air is complete.

FIGS. 32 and 33 illustrate yet another application of the swirl generator of the present invention. In this example, a swirl-augmented combined cycle engine 2400 is illustrated to include a core turbojet engine 2402, a plurality of ramjet engines 2404 that surround the core turbojet engine 2402 and a flow controller 2406. The core turbojet engine 2402 conventionally includes a low pressure compressor 2410, a high pressure compressor 2412, combustors 2414, an air bypass 2416 and a high pressure turbine 2418. As those skilled in the art will appreciate, the core turbojet engine 2402 may optionally include an afterburner 2402 a and a variable area nozzle 2402 b. Although the afterburner 2402 a is illustrated to be a conventional afterburner having a fuel spray ring and a concentric V-gutter flameholder ring, those skilled in the art will appreciate the that the afterburner 2402 a may alternatively be configured in the manner illustrated in FIG. 41 and discussed in detail, below.

The ramjet engines 2404 are configured in a manner that is similar to the ramjet missile of FIG. 29 and 30. Briefly, each of the ramjet engines 2404 includes an air inlet 2420, a swirl generator 30 and a ramjet combustor/nozzle 2430.

The flow controller 2406 is coupled to the core turbojet engine 2402 and includes a forward movable element or diverter 2440, which is employed to selectively control the intake of air into the core turbojet engine 2402 and the ramjet engines 2404, and an aft movable element or diverter 2442, which is employed to selectively close off the outlet of the core turbojet engine 2402 and the ramjet engines 2404. In the particular example provided, the forward and aft movable elements 2440 and 2442 are hingedly mounted to the housing 2450 of the core turbojet engine 2402 and pivotable between a first condition (illustrated in broken line), which closes off the air inlet 2420 and combustor/nozzle 2430, respectively, of each ramjet engine 2404, and a second condition (illustrated in solid line), which closes off the intake side of the low pressure compressor 2410 and the outlet of the high pressure turbine 2418, respectively. The forward and aft movable elements 2440 and 2442 may be moved through any of the various conventionally known means, including hydraulic actuators (not shown).

The core turbojet engine 2402 produces all of the propulsive power that is output by the swirl-augmented combined cycle engine 2400 from zero velocity through a predetermined transition-in velocity of, for example, at about Mach 2. Therefore, the forward and aft movable elements 2440 and 2442 are maintained in the first condition at speeds below the predetermined transition-in velocity.

At the predetermined transition-in velocity, the ramjet engines 2404 are activated to provide additional thrust. Flow entering the ramjet engines 2404 is subjected to additional ramjet compression 2460 and is further compressed to subsonic speeds in a transfer duct/shock isolator 2462. The air flow enters the swirl generator 30 and is converted into a highly turbulent flowfield into which fuel is injected in the manner described above. As the core turbojet engine 2402 is also producing thrust, the forward and aft movable elements 2440 and 2442 are maintained in a position between the first and second conditions. Additionally or alternatively, thrust augmentation by one or more of the ramjet engines 2404 may be accomplished via a bleed-burn process wherein air is tapped-off the high pressure compressor 2412, routed to the ramjet engines 2404 and burnt with added fuel (up to stoichiometric conditions) in the ramjet engines 2404. When the bleed-burn process is to be initiated, a valve 2463 is opened to permit air to flow through a duct 2464 into the ramjet 2404. Those skilled in the art will appreciate that a plurality of guide vanes 96 b, similar to guide vanes 96 discussed above, may be employed to provide a smooth, low loss entry of the bleed air into the ramjet engine 2404.

When air speed reaches a predetermined ramjet takeover velocity of, for example, about Mach 3 to about Mach 4, the forward and aft movable elements 2440 and 2442 are positioned in the second condition and thrust production shifts entirely to the ramjet engines 2404. The swirl ramjets 20 provide the required thrust from about Mach 3 to the cruise condition at speeds of up to approximately Mach 6.

As those skilled in the art will appreciate, the position of the forward and aft movable elements 2440 and 2442 may be controlled in response to flow sensors (not shown) in a selective manner to thereby affect the amount of air that is directed to the core turbojet engine 2402 and the ramjet engines 2404. Alternatively, the forward and aft movable elements 2440 and 2442 may be controlled such that they move continuously from the first condition to the second condition at a predetermined rate upon the sensing of an air speed equivalent to the predetermined transition-in velocity or other event which would prompt the transition from one propulsion mode to another. By incorporating the swirl generator invention into the ramjet engine, this combined cycle engine concept can significantly reduce weight and size, yet provide high performance of the combustor and swirl ramjet concepts compared to other known turbine-based combined cycle engines using conventional ramjet systems. It also reduces the ramjet hot section length, thus reducing cooling requirements.

One unique feature of this concept is the aft-end valving that is accomplished by the aft movable element 2442. The shape of the aft movable element 2442 is such that when positioned in the second condition, the aft movable element 2442 seals-off the aft end of the core turbojet engine 2402 and extends rearwardly in a manner wherein the surface 2442 a of the aft movable element 2442 defines one of the surfaces of the nozzle portion of the ramjet combustor/nozzle 2430. Conversely, when the aft movable element 2442 is positioned in the first position, its surface 2442 b defines a portion of an expansion nozzle for the core turbojet engine 2402.

Another swirl-augmented combined cycle engine 2400 a is illustrated in FIGS. 34 and 35. The swirl-augmented combined cycle engine 2400 a is similar to the swirl-augmented combined cycle engine 2400 of FIGS. 32 and 33 in that it employs turbojet engines 2402, ramjet engines 2404 a and a flow controller 2406 a, which selectively controls the input of air to the turbojet engine 2402 and the ramjet engine 2404 a. The swirl-augmented combined cycle engine 2400 a, however, is globally rectangular, segregated into a plurality of engine cells 2500, with each engine cell 2500 including a turbojet engine 2402, a ramjet engine 2404 a and a flow controller 2406 a. Operation of each engine cell 2500 is identical to the operation of the swirl-augmented combined cycle engine 2400 of FIGS. 32 and 33 and as such, need not be described in detail. The plurality of engine cells 2500 are operated in a manner such that propulsion is regulated between the turbojet engines 2402 and the ramjet engines 2404 a in a uniform manner across the engine cells 2500 (i.e., transition from turbojet propulsion to ramjet propulsion is substantially simultaneous across all of the engine cells 2500).

A third swirl-augmented combined cycle engine 2400 b is illustrated in FIG. 35A. The swirl-augmented combined cycle engine 2400 b is a variant of the two previously described swirl-augmented combined cycle engine configurations in FIGS. 32 and 34, with one major difference being that the core engine 2402 b is coaxial with the ramjet engine 2404 b. In the example provided, the core engine 2402 b is a conventional gas turbine engine that features a co-axial afterburner 5020 having a swirl generator 30 r, thus replacing the long diffuser cone, one or more fuel spray rings and one or more concentric V-gutter flameholder rings.

The key features of the swirl generator include a variable angle swirl vane pack 206 r, a centerbody assembly 236 r, which includes a collapsible centerbody cone 5022, a burner 5024, which employs a quarl step 5026, an array of first fuel injectors 5028, which are integrated (e.g., embedded) into the base of the vanes 402 r of the swirl vane pack 206 r for main afterburning, an array of second fuel injectors 5030 (illustrated as injection orifices in the particular example shown) that are located on the aft centerbody assembly 244 r for ignition and piloting the afterburning, one or more igniters (not shown) that are located in the aft end or base of the aft centerbody assembly 244 r and/or in the recess of the quarl step 5026 and a variable area nozzle 5010.

The swirl-augmented combined cycle engine 2400 b operates in a similar manner to those combined cycle engine concepts already described and therefore will not be discussed in detail other than to note that the swirl generator 30 r is employed for both the turbojet and ramjet engine cycles and therefore allows for a reduction in the overall length of the turbojet.

In operation, the combined cycle engine 2400 b employs a translating spike 5050 and an inner turbojet cowl 5060. The translating spike 5050 is employed to control the shock-on-lip condition during supersonic speeds for both turbojet and ramjet operation, while the translating turbojet cowl 5060 cooperates with the translating spike 5050 to control the flow split during turbojet and ramjet mode operation (i.e., the portion of the intake air flow that is directed to the core turbojet engine 2402 b and the ramjet engine 2404 b). Although the forward portion of the translating spike 5050 is shown with a single cone angle, those skilled in the art will recognize that multiple cone angles may be employed to approximate an isentropic spike to thereby minimize total pressure losses generated by inlet shocks.

During the take-over of propulsion by the ramjet engine 2404 b, the translating spike 5050 is moved forwardly toward the inner turbojet cowl 5060 so as to close off the turbojet air intake 5084 and shift the air flow to the outer cowl inlet 5080 and into the air transfer duct/isolator 5070. The valve 5040 is positioned in the outer air transfer duct/isolator 5070 in an open condition to allow air to flow directed into the swirl generator 30 r and through the nozzle 5010. In this condition, the upstream portion of the turbine air bypass duct 5090 is closed off to prevent feedback and resonance using the same valve 5040 or an alternate valve 5045. Those skilled in the art will appreciate that the valve 5040 is not required if the core turbojet engine 2402 b lacks an air bypass duct 5090. The variable area nozzle 5010 is also repositioned during mode transition (i.e., the transition from turbojet propulsion to ramjet propulsion).

This swirl-augmented combined cycle engine 2400 b allows very compact engine packaging to reduce weight and length, yet provides high performance compared to known turbine based combined cycle engines that utilize conventional gas turbine afterburners and ramjet systems. The swirl generator 30 r/nozzle, which serves as the afterburner for the gas turbine engine and ramjet engine, also reduces the traditional ramjet engine hot section length and in turn reduces cooling requirements to maintain structural and thermal integrity of the hardware.

FIGS. 36 and 37 illustrate yet another application of the swirl generator of the present invention. In this embodiment, a rocket-based combined cycle engine 4000 is illustrated to include a housing 4002, one or more rocket engines 4004, a ramjet engine 4006 and a nozzle 4008. The housing 4002 houses the rocket engines 4004 and the ramjet engine 4006 and defines an air inlet 4020. The rocket engines 4004, which may be coupled to the housing 4002 such that they are located in the dump step 4022 (or quarl surface) and/or inside the aft end of the aft centerbody assembly 244 p, may employ a liquid, slurry or solid fuel, depending upon the application and considerations for the altitude, range and speed that are mandated by the mission. The ramjet engine 4006 includes a swirl generator 30 p which is generally similar to the swirl generator 30, except for the aforementioned rocket engine 4004 that is mounted to aft centerbody assembly 244 p.

The rocket engines 4004 provide low speed thrust and additionally serve to pump air into the air inlet 4020. Air flowing through the air inlet 4020 is converted into a highly turbulent flow field into which fuel is injected and mixed (via the swirl generator 30 p) in the manner described in detail above. In this regard, the air flowing through the air inlet is employed in an afterburning operation by the ramjet engine 4006 to augment the thrust that is generated by the rocket engines 4004 at all speeds. The pumping action is a result of the momentum transfer from the high velocity rocket exhaust and the entrained ambient air. The momentum transfer is the result of the turbulent exchange through the shear layers separating the exhaust of each rocket engine 4004 and the entrained air. Alternatively, the rocket engines 4004 may be used to rapidly accelerate the rocket-based combined cycle engine 4000 above a predetermined speed threshold after which propulsion is transitioned to the ramjet engine 4006.

Those skilled in the art will appreciate that a forward flap or diverter (not shown) may be employed to close-off the air inlet 4020 during low speed operation of the rocket-based combined cycle engine 4000 to effect pure rocket thrust generation up to a predetermined speed threshold after which propulsion is transitioned to the ramjet engine 4006. This mode of operation produces higher ramjet combustor pressures and associated thrust without the occurrence of backflow through the air inlet 4020 at lower speeds.

For optimum thrust in any mode of operation, the nozzle 4008 preferably includes a variable area throat 4030. The variable area throat 4030 is selectably configured to match the flow rate and back-pressure requirements of the rocket-based combined cycle engine 4000 for maximum and efficient thrust generation. Those skilled in the art will appreciate, however, that other nozzle throat concepts may be employed in the alternative, including consumable throat inserts, frangible throat inserts and ejectable throat inserts.

With additional reference to FIGS. 38 and 39, an exemplary variable area throat 4030 is illustrated to include a plurality of throat closure elements 4032 that are rotatable in the housing 4002 through an angle of about 90° between an open position, which is illustrated in FIGS. 37 and 38, and a closed position, which is illustrated in FIG. 39. Although only six elements are shown, those skilled in the art will appreciate that the number of throat closure elements 4032 may be varied to coordinate with the particular upstream geometry of the a rocket-based combined cycle engine 4000. In the particular example illustrated, gaps 4034 are aligned to the six upstream rocket engines 4004 in the low speed mode to minimize erosion of the throat closure elements 4032 during the operation of the rocket engines 4004. When the throat closure elements 4032 are positioned in the closed position during the operation of the ramjet engine 4006, the area of the throat is substantially reduced.

In the embodiment illustrated, the throat closure elements 4032 rotate about an axis on a plane or facet on the inner surface of the housing 4002. The operational mechanism for rotating the throat closure elements 4032 may be housed in the housing 4002 or mounted on an adjacent structure (e.g., fins) to which the rocket-based combined cycle engine 4000 is mounted. Preferably, the throat closure elements 4032 are operated in opposed pairs so as to minimize rotational torques on the rocket-based combined cycle engine 4000 when the throat closure elements 4032 are moved between the open and closed positions.

When the throat closure elements 4032 are positioned in the closed position, the throat closure elements 4032 cooperate to provide the throat with an approximately circular shape; the extent to which the throat is circular is dependent upon the number of throat closure elements 4032 that are employed and whether or not the edges of the throat closure elements 4032 are contoured.

During the operation of the rocket-based combined cycle engine 4000, compressed air enters the air inlet 4020 and is directed to the swirl generator 30 p. The air inlet 4020 also functions as an isolator at supersonic flight speeds where the air is further compressed and brought to subsonic speeds prior to being directed into the swirl generator 30 p. The length of the air inlet duct 4020 is dictated by packaging requirements including fuel, propellants, warhead type (for missile applications), plumbing, controller(s), actuators and batteries. The minimum length of the air inlet duct 4020 is dictated by the isolation requirements that are necessitated during supersonic flight speeds.

During the combined operation of the rocket engines 4004 and the ramjet engine 4006, the swirl generator 30 p functions to augment core flow mixing where flame stabilization is achieved in the recirculation zones in the backward facing areas between the exhaust nozzles 4004 a of the rocket engines 4004 that are mounted in the outer step and/or an annular lip region at the end of the aft bluff boat-tail 304 p. Special igniters are not required since the hot rocket exhaust will serve to ignite the fuel/air mixture for afterburning and ramjet operation. In addition, fuel and/or rocket propellant may be continuously bled through the otherwise idle rocket engines to help cool them and to prevent back flow of the ramjet's hot combustion by-products. The total engine flow passes through the variable area throat 4030 and is expelled to the atmosphere. Although the primary application anticipated for this technology is missile propulsion, those skilled in the art will appreciate that the rocket-based combined cycle engine 4000 may also be employed for aircraft propulsion where reduced weight and complexity would be desired.

Another application of the swirl generator of the present invention is shown in FIG. 41, which illustrates the retrofitting of a compact swirl afterburner 5020 for a conventional afterburner 5000 of the military gas turbine engine 5002 that is shown in FIG. 40. Briefly, the military gas turbine engine 5002, which may be a turbojet engine or a turbofan engine, includes a coaxially-mounted afterburner 5000 having a diffuser tailcone 5004 with one or more fuel spray rings 5006, one or more concentric V-gutter flameholder rings 5008 and a variable area nozzle 5010. The variable area nozzle 5010 is fully opened for afterburning operation and is reduced for non-afterburning operation.

Returning to FIG. 41, retrofit of the military gas turbine engine 5002 entails the substitution of afterburner 5020 for the conventional afterburner 5000 (FIG. 40). The afterburner 5020 includes swirl generator 30 r having a variable-angle swirl vane pack 206 r, a centerbody assembly 236 r, which has a collapsible centerbody cone 5022, an expanding burner 5024, which has a quarl step 5026, an array of fuel injectors 5028 that are embedded into the base of the vanes 402 r for main afterburning, an array of circumferentially spaced apart fuel injectors 5030 (illustrated as injection orifices in the particular example illustrated) located on the aft centerbody assembly 244 r for ignition and piloting the afterburning, and one or more igniters (not shown) that are located in the aft end or base of the aft centerbody assembly 244 r and/or in the recess of the quarl step 5026.

In order to maximize the benefits of the swirl augmentation provided by the swirl generator 30 r, attachment of the swirl generator 30 r should be as close as possible to the turbine exit plane. Accordingly, the length of the conical diffuser (i.e., the collapsible centerbody cone 5022) can be shortened relative to the embodiment of FIG. 40.

The hot gases, which consist mostly of air exiting the turbine and relatively cold air from the bypass fan of the main engine, enter the swirl generator 30 r where they are swirled and the streams are mixed to form a highly turbulent, three-dimensional flowfield. The fuel that is injected into this high shear stress-laden swirling flow is rapidly atomized and mixed. Atomization and mixing are controlled by a novel design of the swirl generator 30 r.

The swirling mixture of the afterburner fuel, hot turbine gases and colder bypass fan air are slowed down across the quarl step 5026 as the flow enters the combustor 5024, and creates a central recirculation zone and an outer recirculation zone similar to the central recirculation zone 610 and outer recirculation zone 640 of FIG. 13. The central recirculation zone is governed by the combined effects of the swirl strength (a characteristic of the variable-angle swirl vane pack 206 r) and the blunt aft end of the aft centerbody assembly 244 r. The outer recirculation zone is created by separation of the fuel/air mixture as it flows over the quarl step 5026. Combustion in the afterburner 5020 is very robust, stable and highly efficient such that the energetic, high temperature by-products of the combustion event are expanded through the variable area nozzle 5010 to provide high levels of thrust.

During non-afterburning operation, the swirl generator 30 r serves as a channel between the main engine and the variable area nozzle 5010. To avoid significant pressure losses due to the presence of the vanes 402 r, the variable-angle swirl vane pack 206 r is controlled such that the angle of the vanes 402 r is changed to 0° so as to remove the swirl from the flow and thereby maintain the axial character of exhaust flow. Vanes 402 r having a flat profile are presently preferred. As an alternative to the variable angle swirl vane pack 206 r, a two-position swirl vane pack (not shown) may also be employed. Also during non-afterburning operation, the collapsible centerbody cone 5022 that is attached to the aft centerbody 244 r is extended to create a flowfield with relatively greater aerodynamic efficiency and relatively lower pressure losses.

With reference to FIG. 41A, a two-position, retractable centerbody aft-cone 5022B is employed for mode transition. The retractable centerbody aft-cone 5022B is fully extended rearward, as is depicted in the portion of the figure above the centerline, during non-afterburning operation to maintain low pressure loss by keeping the flow attached to the aft cone, and is fully retracted forward, as is depicted in the portion of the figure below the centerline, during afterburning operation to provide flow separation with a robust central recirculation zone that is required for flameholding and rapid flame spreading. The housing of the centerbody 5022A has an annular thickness (i.e., defines an annular wall) that contains the manifolds for the fuel injection sites 5028. Various means may be employed to actuate (i.e., extend and retract) the retractable centerbody aft-cone 5022B, such as internal hydraulic or pneumatic actuators that may, for example, be fuel driven. As another alternative, the retractable centerbody aft-cone 5022B may be actuated via a conventional and well-known jackscrew type actuator. Support and guidance for the retractable centerbody aft-cone 5022B may be provided by the actuator and the housing of the centerbody 5022A. In the retracted position, although a portion of the housing of the centerbody 5022A is exposed when the retractable centerbody aft-cone 5022B is retracted, the fuel flowing through it to the fuel injection sites 5028 functions as a coolant which cools the housing of the centerbody 5022A.

Because the afterburner 5020 is able to operate at flamespreading velocities that are over four times greater than those attainable by the conventional afterburner 5000, combustion in the afterburner 5020 is completed in a considerably shorter distance, and therefore it becomes compact, lighter in weight and more fuel efficient than the conventional afterburners. Additionally, no aero-intrusive instream flameholders 5008 of FIG. 40 are required, because the flame stabilization and propagation processes are controlled by the aerodynamics of the swirling flowfield. No aero-intrusive fuel injection rings 5006 of FIG. 40 are needed either, because the fuel is injected from the trailing edges 5028 of the swirl vanes and the centerbody 5030. The long tail cone 5004 is eliminated and replaced with a short centerbody 236 r having an extendable and retractable tapered cone 5022, and variable angle swirl vanes 206 r to accommodate afterburning and non-afterburning modes of operation in order to maximize thrust output at all operating conditions.

The thermal control of the swirl-augmented afterburner could be performed using known combustor technology methods. These techniques would include: fuel scheduling to minimize heat flux to the walls; bypass air cooling, bleed, and film cooling to reduce the temperature of wall materials; and the use of high temperature alloys and materials to withstand the afterburner operating temperatures. Ceramic matrix and other high temperature/refractory materials can be used in localized zones. In addition, fuel injector faces and possibly portions of the exposed centerbody surfaces may be fuel cooled. A specific design would utilize multiple approaches to minimize thermal issues and provide increased service life at the lowest cost.

While the invention has been described in the specification and illustrated in the drawings with reference to a preferred embodiment, it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention as defined in the claims. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiment illustrated by the drawings and described in the specification as the best mode presently contemplated for carrying put this invention, but that the invention will include any embodiments falling within the foregoing description and the appended claims. 

1. An apparatus comprising: at least one core turbojet engine, each core turbojet engine having a low pressure compressor, a high pressure compressor, a combustor, and a turbine; at least one ramjet engine coupled to the at least one core turbojet engine, each ramjet engine having an air inlet, a swirl generator, a ramjet combustor and a nozzle, the swirl generator being coupled to the air inlet and being operable for converting an oxidizer flow into a three-dimensional flowfield that includes a tangential velocity component, the swirl generator including a flow defining means and a fueling means, the flow defining means being operable for effecting both an outer recirculation zone and a central recirculation zone in the ramjet combustor, the outer recirculation zone being toroidal in shape, the central recirculation zone being disposed inwardly of the outer recirculation zone, the fueling means being operable for fueling the outer and central recirculation zones, wherein heat and combustion by-products produced during combustion in the ramjet combustor are carried upstream by the outer and central recirculation zones where the heat and combustion by-products are employed to continuously ignite a combustible fuel/oxidizer mixture in a shear layer adjacent each of the outer and central recirculation zones, which accelerate propagation of the flame into and throughout the core flowfield; and a flow controller coupled to the at least one core turbojet engine and the at least one ramjet engine, the flow controller having at least one forward moveable element and at least one aft moveable element, the at least one forward moveable element being movable between a first position, in which the air intake of the at least one ramjet engine is closed and atmospheric air is permitted to flow into the low pressure compressor of the at least one core turbojet engine, and a second position in which atmospheric air is not permitted to flow into the lower pressure compressor of the at least one core turbojet engine and the air intake of the at least one ramjet engine is open, the at least one aft moveable element being movable between a first position, in which an exit of the nozzle is closed and an exit of the at least one core turbojet engine is open to permit exhaust gas to flow therethrough, and a second position in which the exit of the nozzle is open and the exit of the at least one core turbojet engine is closed.
 2. The apparatus of claim 1, wherein the at least one core turbojet engine includes an air bypass that is selectively operable to input a flow of air into the at least one ramjet engine at a location aft of the air intake.
 3. The apparatus of claim 2, wherein a first end of the air bypass is located between the high pressure compressor and the combustor.
 4. The apparatus of claim 1, wherein the flow defining means includes a swirl vane pack having a plurality of vanes and wherein the vanes are configured to provide the swirl vane pack with a swirl number that is less than about 2.0.
 5. The apparatus of claim 4, wherein the swirl number of the swirl vane pack is about 0.4 to about 1.2.
 6. The apparatus of claim 4, wherein the swirl vanes are mounted to a centerbody assembly.
 7. The apparatus of claim 6, wherein the fueling means includes a plurality of fuel injectors that are coupled to at least one of the air inlet, the centerbody assembly and the swirl vanes.
 8. The apparatus of claim 7, wherein at least a portion of the fuel injectors are positioned to inject the fuel downstream of the swirl vane pack.
 9. The apparatus of claim 7, wherein each swirl vane includes a trailing edge and a lateral surface and wherein the fuel injection sites on a given swirl vane are formed into at least one of the trailing edge and the lateral surface.
 10. The apparatus of claim 6, wherein the centerbody assembly terminates at an aft bluff boat-tail body, the aft bluff boat-tail body being configured to augment flow separation of the exhaust flow.
 11. The apparatus of claim 1, wherein the fueling means includes a plurality of groups of fuel injectors for injecting the fuel directly into the oxidizer flow for at least partially mixing the fuel and oxidizer prior to entering the ramjet combustor.
 12. A method comprising: providing a device having a core turbojet engine, a ramjet engine and a flow controller, the flow controller having a forward moveable element and an aft moveable element the forward moveable element being movable between a first position, in which an air intake of the ramjet engine is closed and atmospheric air is permitted to flow into the core turbojet engine, and a second position in which atmospheric air is permitted to flow only into the ramjet engine, the aft moveable element being movable between a first position, in which an exit of the ramjet is closed and an exit of the core turbojet engine is open to permit exhaust gas to flow therethrough, and a second position in which the exit of the nozzle is open and the exit of the core turbojet engine is closed; positioning each of the forward and aft moveable elements in their respective first positions if a velocity of the device is less than a first transition velocity and operating only the core turbojet engine; and operating both the core turbojet engine and the ramjet engine if the velocity of the device is not less than the first transition velocity and is less than a second transition velocity.
 13. The method of claim 12, wherein the forward and aft moveable elements are positioned between their respective first and second positions when both the core turbojet engine and the ramjet engine are operating.
 14. The method of claim 12, wherein the core turbojet engine includes a high pressure compressor, a combustor and an air bypass, the air bypass including a first end, which is disposed between the high pressure compressor and the combustor, and a second end, which is coupled to the ramjet engine between an air intake and a swirl generator, and wherein a portion of an airflow exiting the high pressure compressor is diverted through the air bypass to the ramjet engine when both the core turbojet engine and the ramjet engine are operating.
 15. The method of claim 12, further comprising operating only the ramjet engine if the velocity of the device is not less than a second transition velocity.
 16. The method of claim 15, wherein the forward and aft moveable elements are positioned in their respective second positions when only the ramjet engine is operating.
 17. The method of claim 15, wherein the second transition velocity is about Mach 3 to about Mach
 4. 18. The method of claim 12, wherein the first transition velocity is about Mach
 2. 19. An apparatus for producing propulsive thrust comprising: a core turbojet engine having a low pressure compressor, a high pressure compressor, a combustor, and a turbine; a ramjet engine having a swirl generator and a ramjet combustor, the ramjet combustor having a combustor inlet, the swirl generator having an inlet housing, a swirl vane pack, a centerbody assembly, and a plurality of fuel injectors, the inlet housing being coupled to the combustor inlet and defining a hollow interior volume therein, the inlet housing serving as a conduit through which an oxidizer flow is conducted, the hollow interior volume intersecting the combustor inlet at a dump step wherein the inlet housing has an inner dimension that is smaller than that of the ramjet combustor, the swirl vane pack being disposed within the hollow interior volume and having a plurality of vanes, the vanes cooperating to change the velocity of the oxidizer flow so that the velocity includes a substantial tangential velocity component, the centerbody assembly being disposed in the hollow interior volume and being coupled to the swirl vane pack so as to extend rearwardly therefrom, and the plurality of fuel injectors being coupled to at least one of the inlet housing, the swirl vane pack and the centerbody assembly and dispensing a fuel therefrom directly into the oxidizer flow for at least partially mixing the fuel and oxidizer prior to entering the ramjet combustor, wherein the swirl generator converts the oxidizer flow received therein into a swirling, three-dimensional flowfield, wherein a first portion of the flowfield flows over the dump step to form an outer recirculation zone, wherein a second portion of the flowfield forms a central recirculation zone that is anchored by an aft end of the centerbody assembly, wherein a first portion of the fuel mixes with the first portion of the flowfield to fuel the outer recirculation zone, wherein a second portion of the fuel mixes with the second portion of the flowfield to fuel the central recirculation zone, and wherein a remaining portion of the fuel enters and fuels a core flow; and a flow controller coupled to the turbojet engine and the ramjet engine, the flow controller having at least one forward moveable element and at least one aft moveable element, the at least one forward moveable element being movable between a first position, in which the air intake of the ramjet engine is closed and atmospheric air is permitted to flow into the low pressure compressor of the core turbojet engine, and a second position in which atmospheric air is not permitted to flow into the lower pressure compressor of the core turbojet engine and the air intake of the ramjet engine is open, the at least one aft moveable element being movable between a first position, in which an exit of the nozzle is closed and an exit of the core turbojet engine is open to permit exhaust gas to flow therethrough, and a second position in which the exit of the nozzle is open and the exit of the core turbojet engine is closed; wherein when a velocity of the apparatus is less than a predetermined first transition velocity, the forward and aft moveable elements are positioned in their respective first positions and the core turbojet engine is operable for producing all of the propulsive thrust, and when the apparatus has a velocity that is not less than the predetermined transition velocity the ramjet engine is operable for producing at least a portion of the propulsive thrust.
 20. The apparatus of claim 19, wherein when a velocity of the apparatus is greater than a second transition velocity, the forward and aft moveable elements are positioned in their respective second positions and the ramjet engine is operable for producing all of the propulsive thrust. 